Airfoil arrangement for a gas turbine engine utilizing a shape memory alloy

ABSTRACT

An airfoil arrangement for a gas turbine engine may include a clearance device using a shape memory alloy movable to provide clearance between an airfoil and one or more other components of the gas turbine engine. The clearance device may be formed as part of a fan blade. The arrangement may be configured to reduce overall weight and dimensions of the gas turbine engine.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims the benefit of Indian Provisional Patent Application No. 202111030863, filed Jul. 9, 2021, the contents of which are incorporated herein in their entirety.

FIELD

The present disclosure relates to an airfoil arrangement for a gas turbine engine.

BACKGROUND

A gas turbine engine generally includes a turbomachine and a rotor assembly. Gas turbine engines, such as turbofan engines, may be used for aircraft propulsion. In the case of a turbofan engine, the rotor assembly may be configured as a fan assembly. The turbomachine may include a spool arrangement. For example, the spool arrangement may include a high pressure, high speed spool and a low pressure, low speed spool. A combustion section of the turbomachine receives pressurized air, which is mixed with fuel and combusted within a combustion chamber to generate combustion gases. The combustion gases are provided to the spool arrangement. For example, the combustion gases may be provided first to a high pressure turbine of the high pressure spool, driving the high pressure spool, and subsequently to a low speed turbine of the low speed spool, driving the low speed spool.

In a turbofan engine, the fan assembly generally includes a fan and a fan casing. The fan generally includes a plurality of airfoils or fan blades extending radially outwardly from a central hub and/or a disk. During certain operations, the fan blades of the fan may rub against or otherwise contact the fan casing, transferring a load to the fan casing and causing damage to the fan blades, the fan casing, or both.

In certain configurations, the fan casing may include a honeycomb material or trench-filler material configured to mitigate load transfer to and through the fan casing. However, this approach is generally costly. Furthermore, this approach may result in larger, heavier, and/or less efficient fan casings.

As such, there is a need for an improved airfoil that reduces load transfer to the fan casing when the rotor blades rub against the fan casing.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a sectional view of a gas turbine engine;

FIG. 2 is a perspective view of a fan assembly for a gas turbine engine;

FIG. 3 is a perspective view of a fan blade for a gas turbine engine;

FIG. 4 is a schematic sectional view of an embodiment of a fan blade for a gas turbine engine;

FIG. 5 is a schematic sectional view of an embodiment of a fan blade for a gas turbine engine;

FIG. 6 is a schematic sectional view of a non-deformed fan blade for a gas turbine engine;

FIG. 7 is a schematic sectional view of a deformed fan blade for a gas turbine engine;

FIG. 8 is a perspective view of a fan blade for a gas turbine engine;

FIG. 9 is a perspective view of a fan blade for a gas turbine engine;

FIG. 10 is a schematic sectional view of an embodiment of a fan blade for a gas turbine engine; and

FIG. 11 is a schematic sectional view of an embodiment of a fan blade for a gas turbine engine.

Other aspects and advantages of the embodiments disclosed herein will become apparent upon consideration of the following detailed description, wherein similar or identical structures may have similar or identical reference numerals.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the invention.

The word “exemplary” is used herein to mean “serving as an example, instance, or illustration.” Any implementation described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other implementations.

As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gas turbine engine or vehicle and refer to the normal operational attitude of the gas turbine engine or vehicle. For example, with regard to a gas turbine engine, forward refers to a position closer to an engine inlet and aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows.

The terms “coupled,” “fixed,” “attached to,” and the like refer to both direct coupling, fixing, or attaching, as well as indirect coupling, fixing, or attaching through one or more intermediate components or features, unless otherwise specified herein.

The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise.

Approximating language, as used herein throughout the specification and claims, is applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about”, “approximately”, and “substantially”, are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value, or the precision of the methods or machines for constructing or manufacturing the components and/or systems. For example, the approximating language may refer to being within a 1, 2, 4, 10, 15, or 20 percent margin.

Here and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

In some embodiments, one or more components of the gas turbine engine described hereinbelow may be manufactured or formed using any suitable process, such as an additive manufacturing process, such as a 3-D printing process. The use of such a process may allow such component to be formed integrally, as a single monolithic component, or as any suitable number of sub-components. In particular, the additive manufacturing process may allow such component to be integrally formed and include a variety of features not possible when using prior manufacturing methods. For example, the additive manufacturing methods described herein enable the production of airfoils, turbine vanes and rotors, compressor vanes and rotors, and/or fan blades. Such components may have unique features, configurations, thicknesses, materials, densities, fluid passageways, headers, and mounting structures that may not have been possible or practical using prior manufacturing methods. Some of these features are described herein.

As used herein, the terms “additively manufactured” or “additive manufacturing techniques or processes” refer generally to manufacturing processes wherein successive layers of material(s) are provided on each other to “build-up,” layer-by-layer, a three-dimensional component. The successive layers generally fuse together to form a monolithic component which may have a variety of integral sub-components. Although additive manufacturing technology is described herein as enabling fabrication of complex objects by building objects point-by-point, layer-by-layer, typically in a vertical direction, other methods of fabrication are possible and within the scope of the present subject matter. For example, although the discussion herein refers to the addition of material to form successive layers, one skilled in the art will appreciate that the methods and structures disclosed herein may be practiced with any additive manufacturing technique or manufacturing technology. For example, embodiments of the present invention may use layer-additive processes, layer-subtractive processes, or hybrid processes.

The gas turbine engine may include a center portion for example including a fan assembly. The fan assembly is generally connected to the spool arrangement. For example, the rotor of the fan assembly may be coupled to the low speed spool. The fan assembly functions to displace air towards the turbomachine. A portion of the air from the fan assembly enters into the turbomachine to move through at least one stage of compression and another portion of air bypasses at least a portion of the turbomachine. Air initially bypassing the turbomachine may enter the turbomachine at a stage further from the fan assembly and air initially entering the turbomachine may bypass further stages of the turbomachine.

Rotation of the rotor serves to accelerate air with the airfoils of the fan towards the turbomachine. The fan casing extends around radially distal ends of the airfoils and may be used to increase performance by reducing blade tip losses. Performance of the fan assembly generally requires maintaining very small clearances between the fan and the fan casing. The fan casing or case may be defined as having an inner surface facing the axis and the center portion rotatable about the axis.

To achieve high performance, tips of the airfoils rotate passing very close to the fan casing. During some events, for example foreign object ingestion into the gas turbine engine, the airfoils may contact the fan casing. In some of these events, the blades may be released from the engine in a scenario known as fan blade out (“FBO”). This scenario poses a threat to the gas turbine engine and its surroundings.

An object of this disclosure is to describe various airfoil arrangements to manage adverse events such as foreign object ingestion. Airfoils of a fan assembly may be configured with a clearance device utilizing a shape memory alloy and movable to provide clearance between an airfoil and one or more other components of the gas turbine engine in these events. The described airfoil arrangements may advantageously provide a reliable, lightweight, and compact configuration to manage these adverse events and to reduce and control FBO events.

To mitigate damage and potentially reduce the need for immediate repair, some fan assemblies are equipped with abradable material. For example, abradable material may be provided on the airfoils and/or an interior of the fan casing. The fan casing may be provided with a trench circumferentially arranged on the interior of the fan casing and generally aligned with the rotation of the fan. Abradable trench filler may be positioned inside the trench and configured to dissipate some of the impact energy from contact of the fan and fan casing. Such configurations can require increased weight and space to effectively manage these events.

A clearance device for an airfoil may be provided, for example to enable a controlled and consistent mode of maintaining clearance between the airfoil and the fan casing that may enable reducing cost and weight and load transfer to a surrounding casing. The airfoil may define a tip portion, a root portion, and an intermediate portion between the tip portion and the root portion. The intermediate portion may include the clearance device or a portion of the clearance device. In some embodiments, the clearance device may extend into some or all of the tip and/or root portions. The root portion may extend between the intermediate portion and an airfoil root along a span S. The tip portion may extend between the intermediate portion and an airfoil tip along the span S. One or more transition portions may be provided. For example, a root transition portion may be provided between the intermediate portion and the root portion. A tip transition portion may be provided between the intermediate portion and the tip portion. The one or more transition portions may be configured to include the intermediate portion and at least one of the root portion and the tip portion. For example, a transition portion may define an overlap between the intermediate portion and another portion. In an embodiment, the tip transition portion defines an overlap along an airfoil thickness of the tip portion and the intermediate portion. In another embodiment, the root transition portion defines an overlap along the airfoil thickness of the root portion and the intermediate portion.

The embodiments generally shown and described herein may enable a controlled and consistent span reduction of the airfoil, such as a fan blade, following a damage event, such as a hard rub against a surrounding fan casing. These embodiments enable the airfoil to fold and/or contract at a desired span of the airfoil, for example to mitigate load transfer to a surrounding casing. These embodiments may further enable the airfoil to fold and/or contract such that excessive or extreme unbalance of the fan may be reduced following a damage event, such as airfoil liberation, foreign object damage (e.g., bird strikes, icing, etc.), or loss of lubrication or damping to a bearing assembly.

Referring now to the Drawings, FIG. 1 is a sectional schematic view of an embodiment of a gas turbine engine 10. The shown embodiment may be used within an aircraft in accordance with aspects of the present subject matter. More particularly, for the embodiment of FIG. 1 , the gas turbine engine is a high-bypass turbofan jet engine, with the gas turbine engine 10 being shown having a longitudinal or axial centerline axis 12 extending therethrough along an axial direction A for reference purposes. The gas turbine engine 10 further defines a radial direction R extended from the centerline 12. Although an exemplary turbofan embodiment is shown, it is anticipated that the present disclosure may be equally applicable to turbomachinery in general, such as an open rotor, a turboshaft, turbojet, or a turboprop configuration, including marine and industrial turbine engines and auxiliary power units.

In general, the gas turbine engine 10 includes a turbomachine (indicated generally by reference character 14) and a fan assembly 16 positioned upstream thereof. The turbomachine 14 may include an outer casing 18. The outer casing 18 may be of a substantially tubular configuration and/or may define an annular inlet 20. In addition, the outer casing 18 may further enclose and support a low pressure (“LP”) compressor 22 for increasing the pressure of the air that enters the turbomachine 14 to a first pressure level. A high pressure (“HP”) compressor may be included. For example, a multi-stage, axial-flow configuration of the HP compressor 24 may receive the pressurized air from the LP compressor 22 and further increase the pressure of such air.

The pressurized air exiting the HP compressor 24 may flow to a combustor 26 within which fuel is injected into the flow of pressurized air, with the resulting mixture being combusted within the combustor 26. Combustion products 60 may be directed from the combustor 26 along a hot gas path 67 of the gas turbine engine 10 to a high pressure (“HP”) turbine 28 for driving the HP compressor 24 via a high pressure (“HP”) shaft or spool 30. The combustion products 60 may also flow to a low pressure (“LP”) turbine 32 for driving the LP compressor 22. The LP turbine 32 may also drive the fan assembly 16. For example, the LP turbine 32 may drive the LP compressor 22 and/or the fan assembly 16 via a low pressure (“LP”) shaft or spool 34. The LP shaft 34 may be generally coaxial with the HP shaft 30. After driving the turbines 28 and 32, the combustion products 60 may be expelled from the turbomachine 14 via an exhaust nozzle 36 which may provide propulsive jet thrust.

The turbine sections of the gas turbine engine 10 may include one or more non-rotatable components. For example, the HP turbine 28 may include a plurality of HP turbine stator vanes 29. The LP turbine may include a plurality of LP turbine stator vanes 33. The HP turbine stator vanes 29 and LP turbine stator vanes may be configured to span the hot gas path 67 in the radial direction R. The HP turbine stator vanes 29 and LP turbine stator vanes may be non-rotatable about the longitudinal centerline axis 12 relative to the outer casing 18 and may be used to support the outer casing 18. The HP turbine stator vanes 29 and LP turbine stator vanes 33 may be tuned or tunable to increase the performance of the gas turbine engine 10 by controlling flow speed and pressure.

The turbine sections of the gas turbine engine 10 may include a plurality of rotor blades. For example, the HP turbine 28 may include a plurality of HP turbine rotor blades 31. The LP turbine 32 may include a plurality of LP turbine rotor blades 35. The HP turbine rotor blades 31 and LP turbine rotor blades 35 may be rotatable relative to the outer casing 18. Relative rotation between the HP turbine stator vanes 29 and the HP turbine rotor blades 31 and between the LP turbine stator vanes 33 and the LP turbine rotor blades 35 may be tuned or tunable. For example, pitches of the HP turbine stator vanes 29, HP turbine rotor blades 31, LP turbine stator vanes 33, and/or LP turbine rotor blades may be adjusted or adjustable to optimize flow through the hot gas path 67.

The performance of the gas turbine engine 10 may be increased by minimizing clearances between components such as vanes and blades and ducting components rotatable relative thereto. For example, minimizing and maintaining a clearance between the turbine rotor blades 31, 35 and the outer casing 18 may be used to increase engine performance. Minimizing and maintaining a clearance between the turbine stator vanes 29, 33 and a rotatable component such as the LP shaft or spool 34 or a component coupled thereto may also be used to increase engine performance.

As shown in FIGS. 1 and 2 , the fan assembly 16 of the gas turbine engine 10 may include a fan rotor 38. For example, a rotatable, axial-flow configuration of the fan rotor 38 may configured to rotate in a first circumferential direction P (opposite a second circumferential direction Q) surrounded or ducted by a fan casing 40. In some embodiments, the LP shaft 34 may be connected directly to a center portion comprising the fan rotor 38 and/or the rotor disk 39, such as in a direct-drive configuration. In some configurations, the LP shaft 34 may be connected to the fan rotor 38 via a speed reduction device 37 such as a reduction gear gearbox in an indirect-drive or geared-drive configuration. The speed reduction device 37 may be included between any suitable shafts/spools within the gas turbine engine 10 as desired or required.

In at least certain exemplary embodiments, that the fan casing 40 may be configured to be supported relative to the turbomachine 14. For example, the fan casing 40 may be supported by a plurality of substantially radially-extending, circumferentially-spaced supports. In an embodiment, a plurality of outlet guide vanes 42 are provided as such supports. As such, the fan casing 40 may enclose the fan rotor 38. The fan casing may include an outer nacelle 21, which may be configured for aerodynamic flight properties such as drag reduction. The fan rotor 38 may be connected to a plurality of fan blades 44. For example, a disk 64 may be provided for coupling to the plurality of fan blades 44. Moreover, a downstream portion 46 of the fan casing 40 may extend over an outer portion of the turbomachine 14 to define an airflow conduit 48. The airflow conduit 48 may be configured as a secondary, or by-pass, airflow conduit. The airflow conduit 48 may be configured to provide additional propulsive jet thrust and/or may be used to effect cooling of the gas turbine engine 10, for example using a fan nozzle 47.

A hub 65 may be provided to cover the disk 64. For example, the hub 65 may be rotatable with the disk 64 and aerodynamically contoured to control airflow through the plurality of fan blades 44.

A pitch change mechanism (“PCM”) 61 may be provided. For example, the PCM 61 may be operatively coupled to the plurality of fan blades 44 and configured to vary the pitch of at least some of the fan blades. Individual fan blades 44 may be rotatable relative to the disk 39 about a pitch axis P and controlled by the PCM 61. The PCM 61 may be configured to vary the pitch of fan blades 44 individually or in unison.

It will be appreciated, however, that the gas turbine engine 10 is provided by way of example only. In other exemplary embodiments, the gas turbine engine 10 may have any other suitable configuration. For example, in other exemplary embodiments, the gas turbine engine 10 may include a fixed-pitch fan, may be configured as a direct drive engine (i.e., may not include the gearbox 37), may include any suitable or desired number or configuration of shafts or spools, compressors, and/or turbines, etc.

Referring still to FIG. 1 , during operation of the gas turbine engine 10, it should be appreciated that an initial airflow 50 may enter the gas turbine engine 10 through an initial inlet 52 of the fan casing 40. The initial airflow 50 may then pass through the fan blades 44. Downstream of the fan blades 44, the initial airflow 50 may be divided. For example, the initial airflow 50 may be split into a first compressed airflow 54 that moves through the airflow conduit 48 and a second compressed airflow 56 that enters the LP compressor 22. The second compressed airflow 56 may travel along a core air flow path 15. The pressure of the second compressed airflow 56 may then be increased as it enters the HP compressor 24 as a third compressed airflow 58. After mixing with fuel and being combusted within the combustor 26, the combustion products 60 may exit the combustor 26 and flow through the HP turbine 28. Thereafter, the combustion products 60 may flow through the LP turbine 32 and exit the exhaust nozzle 36 to provide thrust for the gas turbine engine 10.

Referring now to FIGS. 2 and 3 , exemplary airfoil 62 embodiments are provided in the context of a fan blade 44. FIG. 2 as above depicts componentry of the fan assembly 16. FIG. 3 depicts an individual fan blade 44 for use in a fan assembly 16. Although the illustrated airfoils 62 are shown as part of a fan blade 44, it is understood that the following discussion of an airfoil 62 may be equally applied to another airfoil embodiment, e.g., a stator vane or rotor blade of a compressor 22, 24 and/or turbine 28, 32 (see FIG. 1 ). As shown, each fan blade 44 extends radially outwardly along a span S defining a spanwise direction from an airfoil root 64 to an airfoil tip 66. The span S is defined by a distance in the spanwise direction along a span centerline of the fan blade 44 from the airfoil root 64 to the airfoil tip 66. A pressure side 68 and a suction side 70 of the airfoil 62 extend from a leading edge 72 to a trailing edge 74 of the fan blade 44 and between the airfoil root 64 and airfoil tip 66 along the span S. Further, it should be recognized that the airfoil 62 may define a chordwise direction along a chord C at each point along the span S and extending between the leading edge 72 and the trailing edge 74. The chord C is defined by a distance in the chordwise direction along a chord centerline, the chord centerline generally being orthogonal to the span centerline of the fan blade 44 from the leading edge 72 to the trailing edge 74. Further, the chord C may vary along the span S of the airfoil 62. For instance, in the depicted embodiment, the chord C increases along the span S toward the airfoil tip 66. Though, in other embodiments, the chord C may be approximately constant throughout the span S or may decrease from the airfoil root 64 to the airfoil tip 66.

As shown particularly in FIG. 3 , the airfoil 62 may define a thickness extending between the pressure side 68 and the suction side 70 at each point along the span S in a thickness direction T. In certain embodiments, the thickness may be approximately constant throughout the span S of the airfoil 62. In other embodiments, the airfoil 62 may define a variable thickness between the airfoil root 64 and the airfoil tip 66. For instance, the thickness may generally decrease along the span S toward the airfoil tip 66. Additionally, the airfoil 62 may define an approximately constant thickness along the chord C at each point along the span S. Or, in other embodiments, at least one point along the span S of the airfoil 62 may define a variable thickness along the chord C. For instance, the airfoil 62 may define a maximum thickness at a position along the chord C at each point along the span S.

The airfoil 62 may also include a plurality of surfaces. For example, the airfoil 62 may include a first outer surface, a second outer surface opposite the first outer surface. The first outer surface may be, for example, the pressure side 68 of the airfoil 62 and the second outer surface may be the suction side 70 of the airfoil 62 as shown in FIGS. 2 and 3 . The airfoil 62 may enclose a shell formed at least in part by the first outer surface, the second outer surface, a leading edge 72 surface connecting the first and second outer surfaces, and a trailing edge 74 surface spaced apart from the leading edge surface and connecting the first and second outer surfaces.

The fan blade 44 of FIG. 3 may have a hollow configuration. In another example, the fan blade 44 may include one or more cavities for purposes such as cooling. Alternatively, the fan blade 44 may be of solid construction.

For the embodiment shown, each fan blade 44 includes an integral component having an axial dovetail 76. When mounted within the fan rotor disk 39, as illustrated in FIG. 2 , the dovetail 76 is disposed in a dovetail slot of the fan rotor disk 39, thereby attaching the fan blades 44 as part of the fan rotor 38.

In an embodiment, the airfoil 62 may include at least one composite ply (such as one of the plies 82 shown in FIG. 7 , described below). For instance, the airfoil 62 may be formed at least partially from a ceramic matrix composite. More particularly, in certain embodiments, the airfoil 62 may be formed in part from one or more composite plies 82 configured as ceramic matrix composite prepreg plies.

Composite materials may include, but are not limited to, metal matrix composites (MMCs), polymer matrix composites (PMCs), or ceramic matrix composites (CMCs). Composite materials, such as may be utilized in the airfoil 62, generally include a fibrous reinforcement material embedded in matrix material, such as polymer, ceramic, or metal material. The reinforcement material serves as a load-bearing constituent of the composite material, while the matrix of a composite material serves to bind the fibers together and act as the medium by which an externally applied stress is transmitted and distributed to the fibers.

Exemplary CMC materials may include silicon carbide (SiC), silicon, silica, or alumina matrix materials and combinations thereof. Ceramic fibers may be embedded within the matrix, such as oxidation stable reinforcing fibers including monofilaments like sapphire and silicon carbide (e.g., Textron's SCS-6), as well as rovings and yarn including silicon carbide (e.g., Nippon Carbon's NICALON®, Ube Industries' TYRANNO®, and Dow Corning's SYLRAMIC®), alumina silicates (e.g., Nextel's 440 and 480), and chopped whiskers and fibers (e.g., Nextel's 440 and SAFFIL®), and optionally ceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite). For example, in certain embodiments, bundles of the fibers, which may include a ceramic refractory material coating, are formed as a reinforced tape, such as a unidirectional reinforced tape. A plurality of the tapes may be laid up together (e.g., as the composite plies 82) to form a preform component. The bundles of fibers may be impregnated with a slurry composition prior to forming the preform or after formation of the preform. The preform may then undergo thermal processing, such as a cure or burn-out to yield a high char residue in the preform, and subsequent chemical processing, such as melt-infiltration with silicon, to arrive at a component formed of a CMC material having a desired chemical composition.

The airfoil 62 may include various metallic components. For example, the leading edge 72, trailing edge 74, tip 66, and/or root 64 may be formed of at least one metal layer. For example, reinforcement and/or impact resistant features may be provided in these components of the airfoil 62. These reinforcement and/or impact resistant features may be manufactured from a metal, such as titanium. The titanium may be a titanium alloy such as Ti-6Al-4V. Alternatively, iron alloys, steel, and/or aluminum alloys may be provided. For example, the reinforcement and/or impact resistant features may be manufactured from a stainless steel such as 15-5 PH.

As shown in FIG. 3 , the fan blade 44 may also be provided with one or more cover elements. The embodiment shown provides a trailing edge cover 85 encompassing the trailing edge 74, a leading edge cover 87 encompassing the leading edge 72, and a tip cover 91 encompassing the tip 66. As shown in more detail in FIGS. 4 and 5 , a pressure surface cover 89 encompassing the pressure side 68 and a suction surface cover 90 encompassing the suction side 70 may also be provided. It should be appreciated that each of these cover elements may be included or excluded as required without regard to the other cover elements. The cover elements 89, 90, 91 may be provided on respective components to provide impact resistance against foreign objects. The cover elements 85, 87, 89, 90, 91 may be formed of various materials. For example, the cover elements 85, 87, 89, 90, 91 may be formed of composite materials and/or metallic materials.

In an embodiment, at least one of the cover elements 85, 87, 89, 90, 91 is formed as a SMA element. SMA elements can prevent or reduce vibration and provide impact resistance by providing hysteresis damping (pre-stressed) due to the phase transformation properties and variation of loading conditions of SMA. SMA's large deformation strain capability makes SMA elements particularly suitable for absorbing shock or impact. The SMA embodiments of the cover elements 85, 87, 89, 90, 91 may have a thickness between 0.1 and 2000 microns and may be capable of dissipating vibration or modifying a resonance frequency of the fan blade 44 at ambient room temperatures and operational temperatures. It should be understood that the cover elements 85, 87, 89, 90, 91

The gas turbine engine 10 may include a clearance device 84 configured for movement or deflection. For example, a clearance device 84 may be provided to facilitate movement of an airfoil 62 relative to another component of the gas turbine engine 10 such as the fan casing 40. Such an airfoil 62 may be provided on various components of the gas turbine engine 10. For example, one or more of turbine rotor blades 31, 35; turbine stator vanes 29, 33; and/or fan blades 44 may be provided with such a clearance device 84.

More specifically, referring now to FIG. 4 , a schematic, cross-sectional view of an airfoil 62 is provided having a clearance device 84 included therewith. The clearance device 84 may be manufactured using a shape memory alloy (“SMA”). The clearance device 84 may include one or more SMA components or features or may be entirely formed of SMA.

A shape memory alloy as described with reference to the clearance device 84 may be a metal alloy that experiences a temperature-related, solid-state, micro-structural phase change. This configuration of the clearance device 84 may facilitate a change from one physical shape to another physical shape. The temperature at which a phase change occurs generally is called the critical or transition temperature of the alloy. The clearance device 84 may be constructed of various SMA materials. For example, the clearance device 84 may include a titanium nickel alloy frequently called Nitinol alloy. The clearance device 84 may alternatively or additionally include a high temperature type of SMA. For example, the clearance device 84 may be constructed of alloys of Ru alloyed with Nb or Ta. In an embodiment, high temperature SMA may facilitate the clearance device 84 to develop shape memory transition temperatures ranging between about 20 degrees Celsius and about 1400 degrees Celsius. The transition temperature of the clearance device 84 may be tunable to specific applications. For example, a clearance device 84 employed in a fan blade may benefit from a relatively low transition temperature whereas a clearance device 84 employed in an HP turbine rotor blade 31 may benefit from a relatively high transition temperature.

In some embodiments, clearance device 84 may include the SMA material as a major constituent, in an amount greater than 50 weight percent (“wt. %”) of the clearance device 84. In certain embodiments, the clearance device 84 may be essentially composed of the SMA material (e.g., at least 90 wt. %, such as at least 95 wt. %, such as 100 wt. %). A SMA material is generally an alloy capable of returning to its original shape after being deformed. For instance, SMA materials may define a hysteresis effect where the loading path on a stress-strain graph is distinct from the unloading path on the stress-strain graph. Thus, SMA materials may provide improved hysteresis damping as compared to traditional elastic materials. Further, SMA materials may act as a lightweight, solid-state alternative to traditional actuators. For instance, certain SMA materials may be heated in order to return a deformed SMA to its pre-deformed shape. A SMA material may also provide varying stiffness, in a pre-determined manner, in response to certain ranges of temperatures. The change in stiffness of the shape memory alloy is due to a temperature related, solid state micro-structural phase change that enables the alloy to change from one physical shape to another physical shape. The changes in stiffness of the SMA material may be developed by working and annealing a preform of the alloy at or above a temperature at which the solid state micro-structural phase change of the shape memory alloy occurs. The temperature at which such phase change occurs is generally referred to as the critical temperature or transition temperature of the alloy. In the manufacture of a clearance device 84 comprising SMA (also referred to as an SMA clearance device 84) intended to change stiffness during operation gas turbine engine 10, the clearance device 84 may be formed to have one operative stiffness (e.g., a first stiffness) below a transition temperature and have another stiffness (e.g., a second stiffness) at or above the transition temperature.

Some shape memory alloys used herein are characterized by a temperature-dependent phase change. These phases include a martensite phase and an austenite phase. The martensite phase generally refers to a lower temperature phase. Whereas the austenite phase generally refers to a higher temperature phase. The martensite phase is generally more deformable, while the austenite phase is generally less deformable. When the shape memory alloy is in the martensite phase and is heated to above a certain temperature, the shape memory alloy begins to change into the austenite phase. The temperature at which this phenomenon starts is referred to as the austenite start temperature (As). The temperature at which this phenomenon is completed is called the austenite finish temperature (Af). When the shape memory alloy, which is in the austenite phase, is cooled, it begins to transform into the martensite phase. The temperature at which this transformation starts is referred to as the martensite start temperature (Ms). The temperature at which the transformation to martensite phase is completed is called the martensite finish temperature (Mf). As used herein, the term “transition temperature” without any further qualifiers may refer to any of the martensite transition temperature and austenite transition temperature. Further, “below transition temperature” without the qualifier of “start temperature” or “finish temperature” generally refers to the temperature that is lower than the martensite finish temperature, and the “above transition temperature” without the qualifier of “start temperature” or “finish temperature” generally refers to the temperature that is greater than the austenite finish temperature.

In some embodiments, a clearance device 84 comprising SMA may define a first stiffness at a first temperature and define a second stiffness at a second temperature, wherein the second temperature is different from the first temperature. Further, in some embodiments, one of the first temperature or the second temperature is below the transition temperature and the other one may be at or above the transition temperature. Thus, in some embodiments, the first temperature may be below the transition temperature and the second temperature may be at or above the transition temperature. While in some other embodiments, the first temperature may be at or above the transition temperature and the second temperature may be below the transition temperature. Further, various embodiments of SMA clearance devices 84 described herein may be configured to have different first stiffnesses and different second stiffnesses at the same first and second temperatures.

Non-limiting examples of SMAs that may be suitable for forming various embodiments of the SMA clearance device 84 described herein may include nickel-titanium (NiTi) and other nickel-titanium based alloys such as nickel-titanium hydrogen fluoride (NiTiHf) and nickel-titanium palladium (NiTiPd). However, it should be appreciated that other SMA materials may be equally applicable to the current disclosure. For instance, in certain embodiments, the SMA material may include a nickel-aluminum based alloys, copper-aluminum-nickel alloy, or alloys containing zinc, zirconium, copper, gold, platinum, and/or iron. The alloy composition may be selected to provide the desired stiffness effect for the application such as, but not limited to, damping ability, transformation temperature and strain, the strain hysteresis, yield strength (of martensite and austenite phases), resistance to oxidation and hot corrosion, ability to change shape through repeated cycles, capability to exhibit one-way or two-way shape memory effect, and/or a number of other engineering design criteria. Suitable shape memory alloy compositions that may be employed with the embodiments of present disclosure may include, but are not limited to NiTi, NiTiHf, NiTiPt, NiTiPd, NiTiCu, NiTiNb, NiTiVd, TiNb, CuAlBe, CuZnAl and some ferrous based alloys. In some embodiments, NiTi alloys having transition temperatures between 5° C. and 150° C. are used. NiTi alloys may change from austenite to martensite upon cooling.

Moreover, SMA materials may also display superelasticity. Superelasticity may generally be characterized by recovery of large strains, potentially with some dissipation. For instance, martensite and austenite phases of the SMA material may respond to mechanical stress as well as temperature induced phase transformations. For example, SMAs may be loaded in an austenite phase (i.e. above a certain temperature). As such, the material may begin to transform into the (twinned) martensite phase when a critical stress is reached. Upon continued loading and assuming isothermal conditions, the (twinned) martensite may begin to detwin, allowing the material to undergo plastic deformation. If the unloading happens before plasticity, the martensite may generally transform back to austenite, and the material may recover its original shape by developing a hysteresis.

Referring still to FIG. 4 , it will be appreciated that the airfoil 62 may be incorporated into the fan blade 44 of FIG. 3 . For example, FIG. 4 depicts the airfoil 62 as a component including the aerodynamic working surfaces of the fan blade 44.

As above, FIG. 4 depicts a schematic, cross-sectional view of the fan blade 44 of FIG. 3 , including the clearance device 84. As is depicted, the clearance device 84 is enclosed at least partially within the fan blade 44. The clearance device 84 may be inserted inside the fan blade 44 during manufacturing or assembly of the fan blade 44, or may be embedded in the fan blade 44 during formation of the fan blade 44. In an embodiment, the clearance device 84 is additively manufactured with the fan blade 44.

Referring still to FIG. 4 , the airfoil 62 includes a tip portion 86. The tip portion 86 may be configured to selectively fracture, disintegrate, or otherwise lose mass. For example, the tip portion 86 of the fan blade 44 may be a frangible embodiment of the tip portion configured to preferentially fracture upon contact with the fan casing 40. In this manner, it may be possible to avoid or minimize the use of protective barriers or abradable material, such as trench filler, in the fan casing 40. Specific configurations of the clearance device 84, alone or in combination with a frangible tip portion 86 may facilitate such weight saving configurations. The inner surface of the fan casing 40 may be configured to deform by a lesser degree than fan blades 44 rotatable about the axis A upon contact therebetween.

The tip portion 86 is disposed outward of the root portion 88 in the radial direction R. As depicted in FIG. 4 , the tip portion 86 makes up at least part of the tip 66 of an airfoil 62. In an embodiment, an entire external surface of the tip 66 is composed of a frangible embodiment of the tip portion 86. The frangible tip portion 86 may be adjacent the clearance device 84. In an embodiment, the frangible tip portion 86 is arranged substantially surrounding at least a portion of the clearance device 84. In an embodiment, the frangible tip portion 86 is configured to permanently deform responsive to an external load.

Referring still to FIG. 4 , the clearance device 84 is configured to connect two or more components of the fan blade 44. For example, in the embodiment depicted, the clearance device 84 extends between and connects the root portion 88 to the tip portion 86. The root portion 88 of the airfoil 62 may be configured to selectively maintain structural integrity. The clearance device 84 may be permanently attached to both the root portion 88 and the tip portion 86. For example, the clearance device 84 may be permanently affixed to the root portion 88 and the tip portion 86.

The clearance device 84 may be movable between a baseline position and a deflected position. For example, an airfoil 62 including the clearance device 84 may be disturbed or impacted such that the clearance device 84 is deflected. In this manner, a clearance device 84 may take advantage of advantageous properties of its shape memory alloy construction as described above. An embodiment of an airfoil 62 configured with a clearance device 84 can be shown in FIG. 3 . The deflection of the airfoil 62 may be directly translated to deflection of the clearance device 84 into its deflected position.

The clearance device 84 may be disposed partially within the tip portion 86 or vice versa. For example, a portion of the clearance device 84 may be entirely encased in the tip portion 86 and/or a portion of the tip portion 86 may be entirely encased in the clearance device 84. The clearance device 84 may also be integral with the tip portion 86. For example, the clearance device 84 may be co-molded with the tip portion 86. It should be appreciated that the clearance device 84 and the tip portion 86 may be manufactured in the same process or in distinct processes. In an embodiment, one or both of the clearance device 84 and the tip portion 86 are additively manufactured.

The clearance device 84 may be disposed partially within the root portion 88 or vice versa. For example, a portion of the clearance device 84 may be entirely encased in the root portion 88 and/or a portion of the root portion 88 may be entirely encased in the clearance device 84. The clearance device 84 may also be integral with the root portion 88. For example, the clearance device 84 may be co-molded with the root portion 88. It should be appreciated that the clearance device 84 and the root portion 88 may be manufactured in the same process or in distinct processes. In an embodiment, one or both of the clearance device 84 and the tip root portion 88 are additively manufactured.

FIG. 5 depicts a schematic, cross-sectional view of a further embodiment of the fan blade 44 of FIG. 3 , including the clearance device 84. As shown in FIG. 5 , the clearance device 84 may extend to the tip 66 of the fan blade. The depicted embodiment shows that the clearance device 84 may overlap with both the root portion 88 and the tip portion 86.

Referring to FIGS. 4 and 5 , the clearance device 84 as shown comprises a tip intermediate portion 92 defining overlap at a common radial distance of the tip portion 86 and the clearance device 84. As in FIG. 4 , the tip portion 86, which may include the pressure surface cover 89 and/or the suction surface cover 90, may encompass at least part of the clearance device 84. As shown in FIG. 5 , the clearance device 84 may encompass at least part of the tip portion 86. A portion of the clearance device 84 that overlaps the tip portion 86 may be defined as a first clearance device overlapping portion 96, whereas a portion of the tip portion 86 that overlaps the clearance device 84 may be defined as a tip overlapping portion 98.

The first clearance device overlapping portion 96 of the clearance device 84 extends across the chord C. The clearance device 84 at the first clearance device overlapping portion 96 may be defined by a lesser extent across the chord C by percentage than the clearance device 84 at a transition point Y disposed between the root portion 88 and the tip portion 86. For example, the clearance device 84 may be diminished in extent across the chord C by percentage at the first clearance device overlapping portion 96 relative to at the transition point Y by 1, 2, 5, 10, 15, 20, or 25 percent. In an embodiment, the first clearance device overlapping portion 96 extends across less than seventy-five percent (75%) of the chord C.

Still referring to FIGS. 4 and 5 , the clearance device 84 as shown comprises a root intermediate portion 94 defining overlap at a common radial distance of the root portion 88 and the clearance device 84. The clearance device 84 may encompass at least part of the root portion 88. A portion of the clearance device 84 that overlaps the root portion 88 may be defined as a second clearance device overlapping portion 97, whereas a portion of the root portion 88 that overlaps the clearance device 84 may be defined as a root overlapping portion 99. It should be understood that the root portion 88 may also encompass at least part of the clearance device 84.

The second clearance device overlapping portion 97 of the clearance device 84 extends across the chord C. The clearance device 84 at the second clearance device overlapping portion 97 may be defined by a lesser extent across the chord C by percentage than the clearance device 84 at the transition point. For example, the clearance device 84 may be diminished in extent across the chord C by percentage at the second clearance device overlapping portion 97 relative to at the transition point Y by 1, 2, 5, 10, 15, 20, or 25 percent. In an embodiment, the second clearance device overlapping portion 97 extends across less than seventy-five percent (75%) of the chord C. As shown, the second clearance device overlapping portion 97 may be divided by the root overlapping portion 99. In some embodiments, the extent of the clearance device 84 at the second clearance device overlapping portion 97 across the chord may be less than that of the clearance device 84 at the first clearance device overlapping portion 96, for example by 1, 2, 5, 10, 15, 20, or 25 percent.

As shown in FIGS. 4 and 5 , the tip intermediate portion 92 is disposed radially further than the transition point Y from the axis and the root intermediate portion 94 is disposed radially closer than the transition point to the axis. In an embodiment, the root intermediate portion 94 is disposed radially further from the root 64 than the tip 66.

FIG. 6 depicts a schematic sectional view of a fan blade 44 in a non-deformed state as also shown in the previous figures. As illustrated, the non-deformed state may be defined as a first position wherein a clearance line XYZ is defined along a proximal point X, the transition point Y, and a distal point Z. In the non-deformed state, the clearance line XYZ extends in a substantially straight direction along the proximal point X, the transition point Y, and the distal point Z. The clearance line XYZ may generally align with the span S of the airfoil and/or the radial direction R of the gas turbine engine 10, wherein, of the points defining the clearance line XYZ, the proximal point X is nearest the airfoil root 64 and the distal point Z is nearest the airfoil tip 66. In this example, the positioning of the points defining the clearance line XYZ may be defined relative to each other, as above, or relative to another feature such as the airfoil root 64. For example, the proximal point X may be a first distance from the airfoil root 64, the transition point Y may be a second distance from the airfoil root 64, and the distal point Z may be a third distance from the airfoil root 64, where the first, second, and third distances increase in magnitude in ascending order.

FIG. 7 depicts a schematic sectional view of the fan blade 44 of FIG. 6 in a deformed state. The deformed state may be defined by second position wherein the clearance line XYZ is interrupted at the transition point to define a proximal line XY between the proximal point X and the transition point Y, a distal line YZ between the distal point Z and the transition point Y, and a deflection angle D between the proximal line XY and the distal line YZ. The deflection angle D in the deformed state differs from one hundred and eighty degrees. For example, when the tip portion 86 is deflected in the axial direction A, the deflection angle D may be less than one hundred and eighty (180) degrees, and in some embodiments may be less than 170, 160, 150, or 120 degrees and may be greater than 30 degrees.

The deformed state as shown in FIG. 7 may be defined by local deformation of the clearance device 84. For example, the clearance device 84 may be deformed by compression and/or tension. As shown in FIG. 7 , a portion of the clearance device 84 nearer the pressure side 68 may deform in compression and a portion of the clearance device 84 nearer the suction side 70 may deform in tension.

The clearance device 84 may be configured with a deformation threshold. For example, the clearance device 84 may be tuned to resist folding in certain situations. In an embodiment, the clearance device 84 is tuned to remain in the non-deformed state during crosswind, light rub, and/or small to medium flocking bird impact. The clearance device 84 may be tuned using shape memory alloys to have a relatively small deflection during light impact events below the deformation threshold and to have a relatively large deflection during heavy impact events above the deformation threshold. The deflection of the clearance device 84 may be non-linear relative to impact force, for example having a disproportionately greater deflection per unit force above the deformation threshold. In such a way, the clearance device 84 may deflect within a small range of distances below the deformation threshold and may deflect within a large range of values above the deformation threshold.

The clearance device 84 may be configured for various ranges of deflection. For example, the clearance device 84 may be configured to deflect to a deflection angle D of 160, 150, 140, 130, or 120 degrees. In various embodiments, such deflection angles may be non-permanent applying features of SMA materials in the clearance device 84 such as superelasticity or psuedoelasticity and further features as described below.

The clearance device 84 may be configured to undergo a phase transformation responsive to an adverse event. For example, the clearance device 84 may undergo a phase transformation to a martensite phase responsive to an external load. The clearance device 84 may also be configured to return to an austenite phase responsive to relaxing of the external load. The transition temperature may be a tunable feature to correspond with temperature conditions of the gas turbine engine. For example, during takeoff and high power events, components of the gas turbine engine will experience relatively high temperatures compared to cruising or low power events.

In various embodiments, a transition temperature of the clearance device 84 may be tuned or tunable for the intended use. For example, the transition temperature may be tuned to fall between temperature ranges experienced by the clearance device 84 in low power events and high power events. In an embodiment, the clearance device 84 is configured for use in a fan blade 44 with a transition temperature of at least one hundred degrees Celsius (100° C.). In another embodiment, the clearance device 84 is configured for use in a fan blade 44 with a transition temperature of between twenty and one hundred degrees Celsius (20-100° C.). In other embodiments, the transition temperature may be much higher. For example, in embodiments where the clearance device 84 is configured for use in a HP turbine rotor blade 31, the transition temperature may be between a five hundred and one thousand degrees Celsius (500-1000° C.) or may be greater than one thousand degrees Celsius (1000° C.)

The clearance device 84 may be configured to extend over various dimensions of an airfoil 62. For example, the clearance device 84 may extend over 5%, 10%, 25%, 50%, or 75% of the span S of an airfoil 62. As shown in FIG. 4 , the clearance device 84 may be defined with a clearance device length D. The clearance device may also extend over 25%, 50%, 75%, 90%, or 100% of the chord of an airfoil 62. In an embodiment, a section of the clearance device 84 extends over at least fifty percent (50%) of the span S. In an embodiment, a section of the clearance device 84 extends over at least ninety percent (90%) of the chord C.

A shape memory alloy component may be configured to deform to reduce the span S of an airfoil. For example, the clearance device 84 formed of SMA may be used to reduce the span S through deformation towards the turbomachine 14 as described above. The clearance device 84 may be completely formed of SMA or may be partially formed of SMA.

It should be appreciated that in other embodiments, the fan blade 44 may be configured in any other suitable manner. For example, referring now to FIG. 5 , a cross-sectional schematic view is provided in accordance with another embodiment of the present disclosure. As shown, an intermediate portion 94 is provided in the fan blade 44. In the depicted embodiment, the intermediate portion 94 is disposed between the root portion 88 and the tip portion 86. The intermediate portion 94 may include various materials. For example, the intermediate portion 94 may be formed of a shape memory alloy. The intermediate portion 94 may be configured to facilitate local deformation or elasticity, for example to allow local bending, compression, and/or expansion.

FIG. 8 depicts a perspective view of a fan blade for a gas turbine engine, highlighting the extent of the clearance device 84. As shown in FIG. 8 , the extent of the clearance device 84 along the span S may be defined by the tip intermediate portion 92, the root intermediate portion 94, and an intermediate portion 100 disposed therebetween. The clearance device 84 may be configured to extend over various dimensions of an airfoil 62. For example, the clearance device 84 may extend over 5%, 10%, 25%, 50%, or 75% of the span S of an airfoil 62. As shown in FIG. 4 , the clearance device 84 may be defined with a clearance device length D. The clearance device may also extend over 25%, 50%, 75%, 90%, or 100% of the chord of an airfoil 62. In an embodiment, a section of the clearance device 84 extends over at least fifty percent (50%) of the span S. In an embodiment, a section of the clearance device 84 extends over at least ninety percent (90%) of the chord C.

FIG. 9 depicts a perspective view of a further embodiment of a fan blade 44 employing a clearance device 84. The clearance device 84 may be provided between the root portion 88 and the tip portion 86 in this embodiment or may be omitted in this location. The clearance device 84 may additionally or alternatively include at least one radially outer portion extending to the tip 66 of the fan blade 44. As shown, the clearance device 84 may include a plurality of deformable protrusions 101. The plurality of deformable protrusions 101 may be configured to deform upon impact with another component such as the fan casing 40 as shown in FIG. 1 . To facilitate deformation, the plurality of deformable protrusions 101 may be arranged for deformation through rotational movement of the fan rotor 38 in the first circumferential direction P. For example, one or more of the plurality of deformable protrusions 101 may extend at least partially in the second circumferential direction Q to facilitate folding or deformation. The plurality of deformable protrusions 101 may also extend at least partially in the axial direction A in this manner.

Individual deformable protrusions 101 are shown as including a protrusion outer surface 103, a protrusion leading surface 107, and a protrusion trailing surface 109. The protrusion outer surface 103 is shown as forming part of the tip 66 of the fan blade 44. The protrusion leading surface 107 leads the protrusion trailing surface 109 during operation in the first circumferential direction P. As shown, the deformable protrusions 101 may be angled such that the protrusion leading surface 107 is configured to face generally radially outwards and the protrusion trailing surface 109 is configured to face generally radially inwards. Between individual ones of the plurality of deformable protrusions 101 a protrusion inner surface is provided to connect the protrusion trailing surface 109 of a leading deformable protrusion 101 with the protrusion leading surface 107 of an immediately trailing deformable protrusion 101. Although the plurality of deformable protrusions 101 are depicting in polygonal form, it should be understood that the deformable protrusions 101 may be of any angular or curvilinear configuration.

FIG. 9 depicts the plurality of deformable protrusions 101 for clarity, where the dimensions may be exaggerated. Individual deformable protrusions 101 may have define less than 1 percent of the chord C of the fan blade 44. In various embodiments, individual deformable protrusions define 1, 2, 5, or 10 percent of the chord C of the fan blade 44.

FIG. 9 also depicts that there may be empty space between individual deformable protrusions 101. A gap defined between pairs of deformable protrusions 101 may be substantially equivalent to the chord C defined by an individual deformable protrusion 101. In some embodiments, such a gap may be less than the chord C defined by an individual deformable protrusion 101, for example 10, 20, 30, 40, or 50 percent of the chord C defined by an individual deformable protrusion 101.

FIG. 10 depicts a schematic sectional view of an embodiment of the fan blade 44 of FIG. 9 taken along section 10-10. As shown in FIG. 10 , the plurality of deformable protrusions 101 may be at least partially enclosed on one or more sides, for example by one or more cover elements. As depicted, the suction surface cover 90 and the pressure surface cover 89 are provided to enclose the plurality of deformable protrusions 101. In this way, a configuration of the clearance device 84 employing a plurality of deformable protrusions 101 may maintain an effective seal for driving fluid in the axial direction A and/or may maintain resilience to wear or impact. It should be understood that any combination of cover elements 85, 87, 89, 90, 91 may be provided for these purposes. At least one of the cover elements may be arranged adjacent to the clearance device 84. The cover elements may be formed of any metallic, SMA, and/or composite material. In an embodiment, the suction surface cover 90 and the pressure surface cover 89 are formed of composite materials.

FIG. 11 depicts a schematic sectional view of an embodiment of the fan blade 44 of FIG. 9 taken along section 11-11. FIG. 11 depicts a base portion of the clearance device 84 from which the plurality of deformable protrusions 101 extend. Deformation of the deformable protrusions 101 can draw each towards the base portion in the axial direction, and thus closer to the axis of rotation.

The gas turbine engine 10 may be provided with any of the features and elements as shown and described. The illustrations of the embodiments described herein are intended to provide a general understanding of the structure of the various embodiments. The illustrations are not intended to serve as a complete description of all of the elements and features of apparatus and systems that utilize the structures or methods described herein. Many other embodiments may be apparent to those of skill in the art upon reviewing the disclosure. Other embodiments may be utilized and derived from the disclosure, such that structural and logical substitutions and changes may be made without departing from the scope of the disclosure. Additionally, the illustrations are merely representational and may not be drawn to scale. Certain proportions within the illustrations may be exaggerated, while other proportions may be minimized. Accordingly, the disclosure and the figures are to be regarded as illustrative rather than restrictive.

While this specification contains many specifics, these should not be construed as limitations on the scope of the invention or of what may be claimed, but rather as descriptions of features specific to particular embodiments of the invention. Certain features that are described in this specification in the context of separate embodiments can also be implemented in combination in a single embodiment. Conversely, various features that are described in the context of a single embodiment can also be implemented in multiple embodiments separately or in any suitable sub-combination. Moreover, although features may be described above as acting in certain combinations and even initially claimed as such, one or more features from a claimed combination can in some cases be excised from the combination, and the claimed combination may be directed to a sub-combination or variation of a sub-combination.

One or more embodiments of the disclosure may be referred to herein, individually and/or collectively, by the term “invention” merely for convenience and without intending to voluntarily limit the scope of this application to any particular invention or inventive concept. Moreover, although specific embodiments have been illustrated and described herein, it should be appreciated that any subsequent arrangement designed to achieve the same or similar purpose may be substituted for the specific embodiments shown. This disclosure is intended to cover any and all subsequent adaptations or variations of various embodiments. Combinations of the above embodiments, and other embodiments not specifically described herein, are apparent to those of skill in the art upon reviewing the description.

It is intended that the foregoing detailed description be regarded as illustrative rather than limiting and that it is understood that the following claims including all equivalents are intended to define the scope of the invention. The claims should not be read as limited to the described order or elements unless stated to that effect. Therefore, all embodiments that come within the scope and spirit of the following claims and equivalents thereto are claimed as the invention.

Further aspects are provided by the subject matter of the following clauses:

One aspect provides an airfoil rotatable about an axis in a first circumferential direction to move air in a first axial direction, the airfoil having a root portion; a tip portion; and a clearance device formed of a shape memory alloy and disposed between the root portion and the tip portion, the clearance device movable between: a first position wherein a clearance line is defined along a proximal point, a transition point, and a distal point and extends in a substantially straight direction along the proximal point, the transition point, and the distal point; and a second position wherein the clearance line is interrupted at the transition point to define a proximal line between the proximal point and the transition point, a distal line between the distal point and the transition point, and a deflection angle between the proximal line and the distal line less than one hundred and eighty degrees.

Another aspect provides a fan section for a turbine engine, the fan section having a center portion rotatable about an axis; a case having an internal surface facing the axis; and a plurality of fan blades affixed to the center portion, each of the plurality of fan blades comprising: a root portion formed of a composite material; a tip portion disposed radially outward of the root portion; and an intermediate portion formed of a shape memory alloy and disposed between the root portion and the tip portion.

Another aspect provides an airfoil rotatable about an axis in a first circumferential direction to move air in a first axial direction, the airfoil having a root portion; and a tip portion, the tip portion comprising a clearance device formed of a shape memory alloy, wherein the clearance device is configured to deform toward the axis and in a second circumferential direction opposite the first circumferential direction.

Another aspect provides an airfoil wherein the clearance device extends across at least seventy-five percent (75%) of a chord of the airfoil at the transition point.

Another aspect provides an airfoil wherein the clearance device comprises a tip intermediate portion, the tip intermediate portion disposed radially further than the transition point from the axis, wherein the tip portion comprises a tip overlapping portion, wherein the clearance device comprises a first clearance device overlapping portion, wherein the first clearance device overlapping portion overlaps the tip overlapping portion across the chord of the airfoil, and wherein the first clearance device overlapping portion extends across less than seventy-five percent (75%) of the airfoil.

Another aspect provides an airfoil wherein the clearance device comprises a root intermediate portion, the root intermediate portion disposed radially closer than the transition point to the axis, wherein the root portion comprises a root overlapping portion, wherein the clearance device comprises a second clearance device overlapping portion overlapping the root overlapping portion across the chord of the airfoil, and wherein the second clearance device overlapping portion extends across less than seventy-five percent (75%) of the airfoil.

Another aspect provides an airfoil wherein the clearance device extends across at least ninety percent (90%) of a chord of the airfoil at the transition point.

Another aspect provides an airfoil wherein the transition point is disposed closer to a tip of the airfoil than to a root of the airfoil.

Another aspect provides an airfoil wherein the tip portion is formed of a composite material.

Another aspect provides an airfoil further comprising a cover element, wherein the cover element is positioned over the composite material of the tip portion at one or more of: a leading edge of the airfoil; a trailing edge of the airfoil; a pressure surface of the airfoil; a suction surface of the airfoil; and a tip of the airfoil.

Another aspect provides an airfoil wherein the cover element comprises an edge deformation feature.

Another aspect provides an airfoil wherein the edge deformation feature is formed of a shape memory alloy.

Another aspect provides an airfoil wherein the clearance device comprises a tip portion, and wherein the tip portion is configured to deform toward the axis.

Another aspect provides an airfoil wherein the tip portion comprises a plurality of deformable protrusions.

Another aspect provides an airfoil wherein the plurality of deformable protrusions extends at least partially in a second circumferential direction opposite the first circumferential direction.

Another aspect provides a fan section wherein the intermediate portion is configured to control deformation of a respective one of the plurality of fan blades by a greater degree than the internal surface of the case upon interference of the respective fan blade with the internal surface of the case.

Another aspect provides an airfoil wherein the plurality of deformable protrusions extends at least partially in the second circumferential direction.

Another aspect provides an airfoil wherein the tip portion is formed in part of a composite material.

Another aspect provides an airfoil further comprising a cover element formed of a metallic material, wherein the cover element is positioned over the composite material of the tip portion at one or more of: a leading edge of the airfoil; a trailing edge of the airfoil; a pressure surface of the airfoil; or a suction surface of the airfoil, wherein the cover element is arranged adjacent to the clearance device. 

1. An airfoil rotatable about an axis in a first circumferential direction to move air in a first axial direction, the airfoil comprising: a root portion; a tip portion; and a clearance device formed of a shape memory alloy and disposed between the root portion and the tip portion, the clearance device movable between: a first position wherein a clearance line is defined along a proximal point, a transition point, and a distal point and extends in a substantially straight direction along the proximal point, the transition point, and the distal point; and a second position wherein the clearance line is interrupted at the transition point to define a proximal line between the proximal point and the transition point, a distal line between the distal point and the transition point, and a deflection angle between the proximal line and the distal line less than one hundred and eighty degrees.
 2. The airfoil of claim 1, wherein the clearance device extends across at least seventy-five percent (75%) of a chord of the airfoil at the transition point.
 3. The airfoil of claim 2, wherein the clearance device comprises a tip intermediate portion, the tip intermediate portion disposed radially further than the transition point from the axis, wherein the tip portion comprises a tip overlapping portion, wherein the clearance device comprises a first clearance device overlapping portion, wherein the first clearance device overlapping portion overlaps the tip overlapping portion across the chord of the airfoil, and wherein the first clearance device overlapping portion extends across less than seventy-five percent (75%) of the airfoil.
 4. The airfoil of claim 3, wherein the clearance device comprises a root intermediate portion, the root intermediate portion disposed radially closer than the transition point to the axis, wherein the root portion comprises a root overlapping portion, wherein the clearance device comprises a second clearance device overlapping portion overlapping the root overlapping portion across the chord of the airfoil, and wherein the second clearance device overlapping portion extends across less than seventy-five percent (75%) of the airfoil.
 5. The airfoil of claim 4, the clearance device extends across at least ninety percent (90%) of a chord of the airfoil at the transition point.
 6. The airfoil of claim 1, wherein the transition point is disposed closer to a tip of the airfoil than to a root of the airfoil.
 7. The airfoil of claim 1, wherein the tip portion is formed of a composite material.
 8. The airfoil of claim 7, further comprising a cover element, wherein the cover element is positioned over the composite material of the tip portion at one or more of: a leading edge of the airfoil; a trailing edge of the airfoil; a pressure surface of the airfoil; a suction surface of the airfoil; and a tip of the airfoil.
 9. The airfoil of claim 8, wherein the cover element comprises an edge deformation feature.
 10. The airfoil of claim 9, wherein the edge deformation feature is formed of a shape memory alloy.
 11. The airfoil of claim 1, wherein the clearance device comprises a tip portion, and wherein the tip portion is configured to deform toward the axis.
 12. The airfoil of claim 11, wherein the tip portion comprises a plurality of deformable protrusions.
 13. The airfoil of claim 12, wherein the plurality of deformable protrusions extends at least partially in a second circumferential direction opposite the first circumferential direction.
 14. A fan section for a turbine engine, the fan section comprising: a center portion rotatable about an axis; a case having an internal surface facing the axis; and a plurality of fan blades affixed to the center portion, each of the plurality of fan blades comprising: a root portion formed of a composite material; a tip portion disposed radially outward of the root portion; and an intermediate portion formed of a shape memory alloy and disposed between the root portion and the tip portion.
 15. The fan section of claim 14, wherein the intermediate portion is configured to control deformation of a respective one of the plurality of fan blades by a greater degree than the internal surface of the case upon interference of the respective fan blade with the internal surface of the case.
 16. An airfoil rotatable about an axis in a first circumferential direction to move air in a first axial direction, the airfoil comprising: a root portion; and a tip portion, the tip portion comprising a clearance device formed of a shape memory alloy, wherein the clearance device is configured to deform toward the axis and in a second circumferential direction opposite the first circumferential direction.
 17. The airfoil of claim 16, wherein the clearance device comprises a plurality of deformable protrusions.
 18. The airfoil of claim 17, wherein the plurality of deformable protrusions extends at least partially in the second circumferential direction.
 19. The airfoil of claim 16, wherein the tip portion is formed in part of a composite material.
 20. The airfoil of claim 19, further comprising a cover element formed of a metallic material, wherein the cover element is positioned over the composite material of the tip portion at one or more of: a leading edge of the airfoil; a trailing edge of the airfoil; a pressure surface of the airfoil; or a suction surface of the airfoil, wherein the cover element is arranged adjacent to the clearance device. 